Results
94-15-10: This amendment adopts a new airworthiness directive (AD) that is applicable to certain Aerospatiale Model ATR42-300 and -320 series airplanes. This action requires an inspection to verify that the cables supplying power to the aileron trim are adequately protected and to detect any damage to these cables. This amendment also requires installation of protective sleeves on unprotected cables, and replacement of damaged cables. This amendment is prompted by reports of a short circuit of these cables due to worn protective sleeves. The actions specified in this AD are intended to prevent short circuiting of the cables that supply power to the aileron trim, which could lead to loss of aileron trim control.
93-16-01: 93-16-01 AEROSPATIALE: Amendment 39-8658. Docket 92-NM-218-AD. Applicability: Model ATR72-100 and -200 series airplanes, equipped with main landing gear (MLG) side braces having part number (P/N) D23207000 or D23207000-1, certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent collapse of the landing gear, accomplish the following: (a) Prior to the accumulation of 7,500 landings on the MLG side brace or within the next 100 landings after the effective date of this AD, whichever occurs later, replace any side brace having P/N D23207000 or D23207000-1 with a repaired or modified side brace having P/N D23207000-2, in accordance with Aerospatiale Service Bulletin ATR72-32-1011, dated August 7, 1992. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Manager, Standardization Branch, ANM-113. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The replacement shall be done in accordance with Aerospatiale Service Bulletin ATR72-32-1011, dated August 7, 1992. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Aerospatiale, 316 Route de Bayonne, 31060 Toulouse, Cedex 03, France. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on October 7, 1993.
2009-09-51: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 2009-09-51, which was sent previously to all known U.S. owners and operators of Eurocopter France (Eurocopter) Model EC225LP helicopters by individual letters. This AD requires, before further flight, determining if the "CHIP'' detector light on the instrument panel (Vehicle Monitoring System Screen) previously illuminated. If the "CHIP'' detector light did illuminate and it illuminated because of a metal particle on the magnetic plug of the epicyclic reduction gear module (module) of the main gearbox (MGB), or if you cannot determine from the maintenance records which chip detector caused the "CHIP'' detector light to illuminate or whether the detector light stayed illuminated after the "CHIP'' detector switch was turned to the "CHIP PULSE'' setting, replacing the module with an airworthy module is required before further flight. Also required before further flight is inspectingthe MGB module magnetic chip detector electrical circuit and determining whether the system is functioning properly, including whether the "CHIP'' detector light annunciates on the instrument panel (Vehicle Monitoring System Screen). Finally, this AD requires replacing the module with an airworthy module if the "CHIP'' detector light illuminates, stays illuminated after the "CHIP'' detector switch is turned to the "CHIP PULSE'' setting, and you determine that a metal particle on the module magnetic plug caused that illumination. This amendment is prompted by a mandatory continuing airworthiness information (MCAI) AD issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. EASA notified us of an accident that occurred April 1, 2009 on a Eurocopter Model AS332L2 helicopter and EASA advises that the "cause of the accident seems to be connected with degradation of the epicyclic module of the MGB, the root cause ofwhich is still to be determined.'' The actions specified by this AD are intended to prevent failure of the MGB and subsequent loss of control of the helicopter. DATES: Effective December 28, 2009, to all persons except those persons to whom it was made immediately effective by Emergency AD 2009-09-51, issued on April 17, 2009, which contained the requirements of this amendment. Comments for inclusion in the Rules Docket must be received on or before February 9, 2010.
2006-18-09: The FAA is superseding an existing airworthiness directive (AD), which applies to all BAe Systems (Operations) Limited Model ATP airplanes. That AD currently requires revising the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness (ICA) to incorporate life limits for certain items and inspections to detect fatigue cracking in certain structures; to incorporate new inspections to detect fatigue cracking of certain significant structural items (SSIs); and to revise life limits for certain equipment and various components. This new AD requires revising the ALS of the ICA to include additional and revised inspections of the fuselage. This AD results from the manufacturer review of fatigue test results that identified additional and revised inspections of the fuselage that are necessary in order to ensure the continued structural integrity of the airplane. We are issuing this AD to detect and correct fatigue cracking of certain structural elements, which could result in reduced structural integrity of the airplane and consequent rapid decompression of the airplane.
72-07-07: 72-07-07 DOWTY ROTOL: Amdt. 39-1417. Applies to all propellers with blade root sizes Nos. 20 and 30. These propellers are installed on but not limited to Fairchild F-27 Series; Fairchild Hiller F-227 Series; Armstrong Whitworth Argosy Type AW 650 Series; Vickers Viscount 745 and 810 Series; and Grumman G-159 Series airplanes. Compliance is required as indicated unless already accomplished. To prevent possible fatigue failure of propeller blades or blade bearings as a result of loss of bearing pre-load due to discrepancies in the ground tracks or improper assembly of the propeller blade bearing, accomplish the following: (a) For propellers with any bearing identified by one of the Bearing Identification Numbers listed in subparagraph (3) that has not been last reground by an FAA-approved repair facility subsequent to April 21, 1971, or by the bearing manufacturer, comply with paragraph (c)(1) within the next 400 hours' time in service after the effective date of thisAD, or at the next disassembly of a bearing from a blade for cause, whichever occurs sooner. Thereafter, comply with paragraph (c)(2) at the following intervals: (1) For propellers for which there is a scheduled blade overhaul, at each disassembly of a bearing from a blade for cause, including disassembly at overhaul. (2) For propellers for which there is no scheduled blade overhaul, at each disassembly of a bearing from a blade for cause, and at intervals not to exceed 5,000 hours' time in service on a blade from the last inspection. (3) Bearing Identification Numbers: BRGS. BRGS. BRGS. BRGS. BRGS. BRGS. 11309 10449/R 16726 16331/R/R 18305/R 4696/R/R 16350 9660/R 15071/R 947/R/R 18328/R 13098 13917/R/R 15337 13345/R 17078/R N16672/R 18077/R N15311 1237/R/R/R 2280/R/R 2847/R/R N8225/R N17430 3670/R/R N13981 N15427 16970/R 11781/R N17368 7801/R/R N14011 16498/R N17313 2629/R N13980 N4629/R 15523/R/R 12825/R N17424 N13977 20448/R 15959/R N17829/R N16327 12794/R N11240/R N15257/R N16078/R N16167/R 5901/R/R 10571 N21240N/R 14731/R N14495/R N16490/R 7060/R/R N9730/R N16898/R 17082/R 5756/R 18119 1298/R N16532 16831/R N18287 3626/R/R 21243 15309/R N18236 18081/R 11766/R 16790 3258/R 16512/R N12714 11838/R/R 9180/R 4416/R 20441/R/R N18094 2489/R 15302/R N10966 N11149 14352/R 3054/R 16339/R N3682/R/R 18055 N15574 N11761 1400/R/R 11751/R/R 16295 N13837 16179/R/R N14938 11346/R/R N13821 14163 7049/R N11778/R/R N13822 4786/R/R/R 14165/R 14734 N13833 1989/R/R 12749/R/R 1851/R/R/R/R N16977/R N20911N 315/R/R 13347/R 9048/R 4709/R N9296 4354/R/R N4689/R 14019/R/R N15489/R 20045N/R N15890 13969 N434/R/R 21280N/R/R 13993/R N20340 N16773 14386/R/R 15567 17018 11764/R 10422/R/R N16706 15220/R14364/R 10554/R N16647/R 5801/R 2753/R/R 14213/R N15285 11765 N11358/R/R 13559/R/R 16981/R 15275 12038 N21248N/R 12746/R 20449/R/R 15538 11754/R N18103 12571/R/R N15703 N21048N/R 9668 13853/R N15691 20823/R/R 4131/R/R 11756 N16858 5235 5155/R N7479/R N17408 N5007/R/R N11724/R (b) For all other propellers, comply with the following: (1) For propellers for which there is a scheduled blade overhaul - (i) Comply with paragraph (c)(1) at the next disassembly of a bearing from a blade for cause, including disassembly at overhaul; and (ii) Thereafter comply with paragraph (c)(2) at each disassembly of a bearing from a blade for cause, including disassembly at overhaul. (2) For propellers for which there is no scheduled blade overhaul - (i) Comply with paragraph (c)(1) at the next disassembly of a bearing from a blade for cause or before the accumulation of 5,000 hours'time in service on the propeller from the last inspection, whichever occurs sooner; and (ii) Thereafter comply with paragraph (c)(2) at each disassembly of a bearing from a blade for cause and at intervals not to exceed 5,000 hours' time in service on a blade from the last inspection. (c) Inspect blade bearings in accordance with either subparagraph (1) or (2), as applicable. (1) For the initial inspections required by paragraphs (a) and (b) disassemble the propeller to the stage where the assembly of blades complete with bearings are removed from the hub and inspect each blade bearing in accordance with paragraph (d). (2) For the repetitive inspections required by paragraphs (a) and (b), inspect affected blade bearings in accordance with paragraph (d). (d) Measure and record the bearing torque and the center race spigot mean diameter before and after removal of bearing pre-load; and compute the spigot expansion, in accordance with the procedures specified in Dowty Rotol Service Letter No. SER/SSL/1, dated August 12, 1971, or an FAA-approved equivalent. If the spigot expansion, for a bearing size and blade bearing assembly specified in Columns 1 and 2, respectively, of the following table, is - (1) Equal to or greater than Expansion Value A (Column 4), the bearing assembly, blade, and blade bolt may be returned to service. (2) Less than Expansion Value B (Column 5), remove the bearing assembly, blade, and blade bolt from service and mark them in a manner that will prevent their further use. (3) Less than Expansion Value A, but equal to or greater than Expansion Value B (Column 5), before further operation of the blade, blade bolt, and bearing assembly return the bearing assembly to Dowty Rotol, Inc., Arlington, Virginia, or Dowty-Rotol Ltd., Service Department Cheltenham Road, Gloucester, England, for a load-expansion torque test to determine the remaining service life of the blade and blade bolt. The measurements of bearing torque and the mean spigot diameter before and after removal of the blade bearing preload required by this paragraph must accompany the parts. (Col. 1) Bearing Size (Col. 2) Blade Bearing Assembly (Col. 3) Normal Spigot Expansion (Col. 4) Expansion Value A (Col.5) Expansion Value B RA38965 .0055 RA38965Z to .0050 .0035 RA38965Z-1 .0060 No. 20 RA54100 .0050 RA54100Z to .0040 .0030 RA54100Z-1 .0055 RA57099 .0040 RA57099Z to .0035 .0025 RA57099Z-1 .0045 No. 30 RA22355/1 .0035 RA22355/2 to .0030 .0022 .0040 (e) An inspection performed following any period in service since the last installation of the propeller, may be accepted in lieu of the initial inspection required by paragraph (a) to be performed within the next 400 hours' time in service after the effective date of this AD, if the records of thatinspection are submitted to an FAA Maintenance Inspector and he determines that the inspection included a computation of spigot expansion performed in accordance with the procedures specified in paragraph (d) which established that the spigot expansion values were equal to or greater than the applicable Expansion Value A (Column 4) specified in paragraph (d) for the bearing size and blade bearing assembly. This amendment becomes effective March 29, 1972.
90-01-04: 90-01-04 SIKORSKY AIRCRAFT: Amendment 39-6458. Docket No. 89-ASW-29. Applicability: Sikorsky Model S-61 series helicopters certificated in any category, equipped with part number (P/N) S6110-23300 main rotor head. Compliance: Required within the next 150 hours' time in service after the effective date of this AD, unless already accomplished. To prevent possible failure of the main rotor hub upper plate, which could result in the loss of the helicopter, accomplish the following: (a) Conduct a one-time inspection of the main rotor head, P/N S6110-23300, installed with a bifilar assembly, P/N S6112-23039, as follows: (1) Remove main rotor fairing to provide access to the upper hub arms. (2) Loosen the bifilar attachment bolt (NAS630-84) and nut (EB108) a minimum of two turns, one bolt at a time. (3) Using a 0.010-inch feeler gage or equivalent, insert tip of feeler gage between washer stackup and bifilar lug. Accomplish this check on both sides of lug. (i) If tip of feeler gage does not contact shank of attachment bolt, washer stackup and lug clearance are acceptable. Torque attachment nut to 1,690 inch-pounds. (ii) If tip of feeler gage contacts shank of attachment bolt, reshim bifilar support assembly to provide a 0.000 to 0.004 inch clearance. Torque attachment nut to 1,690 inch-pounds maximum. NOTE: The Sikorsky Model S-61 Maintenance Manual contains shimming instructions. (4) Repeat the requirements of paragraphs (a)(2) and (3) individually on the four remaining hub arms for each bifilar support assembly attachment hardware. (5) Reinstall main rotor fairing. (b) Conduct a one-time visual inspection for cracks in area of the 10 main rotor hub upper plate arm bifilar attachment bolt holes. Conduct a visual inspection for cracks on the outside surfaces of each of the 5 hub arms, adjacent to the 10 bifilar attachment holes. The rotor head does not require disassembly for these inspections.(1) If a crack indication is found by the visual inspection, clean immediate area and reinspect using a dye penetrant or equivalent inspection method. (2) If a crack is verified, remove main rotor hub assembly prior to further flight, and replace with an airworthy component. (c) Upon request, an alternate means of compliance which provides an equivalent level of safety to the requirements of this AD may be used when approved by the Manager, Boston Aircraft Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, 12 New England Executive Park, Burlington, Massachusetts 01803, telephone (617) 273-7118. (d) Upon submission of substantiating data by an owner or operator through an FAA Aviation Safety Inspector, the Manager of the Boston Aircraft Certification Office may adjust the compliance time specified in this AD. NOTE: Sikorsky Aircraft Co. Alert Service Bulletin No. 61B10-46, dated February 7, 1989, pertains to this subject. This amendment (39-6458, AD 90-01-04) becomes effective on February 5, 1990.
94-15-11: This amendment adopts a new airworthiness directive (AD), applicable to certain Fokker Model F-27 series airplanes, that requires the implementation of a corrosion prevention and control program either by accomplishing specific tasks or by revising the maintenance inspection program to include such a program. This amendment is prompted by reports of incidents involving corrosion and fatigue cracking in transport category airplanes that are approaching or have exceeded their economic design goal; these incidents have jeopardized the airworthiness of the affected airplanes. The actions specified by this AD are intended to prevent degradation of the structural capabilities of the airplane due to the problems associated with corrosion.
88-09-03: 88-09-03 LOCKHEED-GEORGIA: Amendment 39-5903. Applies to Model 382 series airplanes, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent failure of the control column and loss of control of the airplane, accomplish the following: A. Within the next 250 hours time-in-service after the effective date of this AD, perform an eddy current (non-destructive) inspection of the left and right control column bases on the forward side of each control column at approximately floor level, as follows: NOTE: The control column base is shown in the Lockheed Hercules Component Overhaul Manual, SMP 850, Item 27-26, figure 1, page 4, as Figure Item No. 235 (see Figure 1, below). The inspection area is around the 1.25-inch diameter hole located on the forward side of the base approximately at the lower end of the control column tube support (Figure Item No. 185). The control column bases are fabricated from either magnesium or aluminum. Earlier airplanes were delivered with AZ916-T6 magnesium alloy sand castings per MIL-M-4204. The magnesium castings are finished with an anodic treatment, a wash primer per MIL-C-8514, and two coats of epoxy primer per LAC 37-722, Type I, followed by an epoxy enamel per LAC 37-722 of appropriate color. The aluminum bases are A356-T6 aluminum castings per MIL-A- 21180, Class 10, with an initial sulfuric acid anodize, and finished with two coats of zinc chromate primer, followed by two coats of Camouflage lacquer per TT-L-190. Fatigue cracking may initiate at the sharp edge formed by thickness transition approximately mid-way up from the bottom of the 1.25-inch diameter hole on the forward side of the base. Cracks will propagate aft around the perimeter of the base along the thickness transition. 1. Prior to performing the eddy current inspection, assemble the following equipment: a. Crack Detector, Magnaflux P/N ED-520, or equivalent. b.Probe, surface, shielded, 1/8-inch diameter, 8 inches long, VM products P/N VM200-8, or equivalent. c. Probe, surface, shielded, right angle, 1/8-inch diameter, 1/2-inch drop, 8 inches long, VM Products P/N VH202BA-1/2-8, or equivalent. d. Cable, 6 foot, Microdot/BAC, Mastercraft Enterprises P/N NMT 7-41-BM, or equivalent. e. Standard, aluminum, as furnished with the Magnaflux instrument, or Lockheed-Georgia Company P/N PDL 457A. 2. Prior to performing the eddy current inspection, prepare the airplane in the following manner: WARNING: Ensure that power is isolated from all systems in the inspection area prior to approaching the inspection area. Failure to comply may result in serious injury to personnel. a. Isolate power from all systems in the inspection area in accordance with applicable maintenance manuals. b. Remove screws retaining Dust Seal Boot (Figure Item No. 20, Item 27-26 of SMP 850) to the cockpit floor. Loosen clamp retaining the Dust Seal Boot to the Support Tube, and push the Dust Seal Boot up approximately 6 - 8 inches and secure to Support Tube temporarily. Repeat for opposite control column installation. 3. Gain access for the inspection inside of the airplane in the cockpit, using normal ingress provisions. 4. Prepare the area to be inspected as follows: a. Remove oil grease, residual, sealant, and other materials by wiping the control column base around the 1.25-inch diameter hole on the forward side of the base with Federal Specification 0-T-620, Type I, Trichloroethane, or commercial equivalent, and dry with clean cotton cloth. WARNING: Federal Specification 0-T-620, Type I, Trichloroethane is toxic to skin, eyes, and respiratory tract. Skin and eye protection is required. Avoid repeated or prolonged contact. Good general ventilation is normally adequate. Failure to comply could result in injury to personnel. b. Visually inspect area to be inspected, looking for rough surface conditions which could adversely affect the eddy current inspection. If surface appears rough, smooth surface using 320 grit sandpaper until a reliable eddy current inspection can be achieved. 5. Set up and calibrate the eddy current instrument as specified in Lockheed Service Manual SMP-515-A or SMP-515-C, Work Card SP-76, Appendix A, for surface scanning on aluminum. The eddy current instrument must be recalibrated when probes are changed. Calibrate on aluminum for inspection of the magnesium casting. 6. Perform the eddy current inspection in the following manner: a. With the control column in the full aft position, surface scan around the entire circumference of the 1.25-inch diameter hole at the front of the control column base casting, using both eddy current probes. Scan in both clockwise and counterclockwise directions. b. Repeat this inspection on the opposite control column base. 7. If any suspected defects in thecontrol column base casting are found during the eddy current inspection, indicated by a sharp meter deflection noted during calibration, mark the suspected defect. a. To confirm suspected defects, proceed to paragraph B., below. b. If no defects are suspected, proceed to paragraph C., below (System Securing). B. To confirm suspected defects, perform a fluorescent penetrant inspection as follows: 1. Prior to performing the fluorescent penetrant inspection, assemble the following equipment: NOTE: All chemicals shall use the "family concept" - they shall all be from the same manufacturer for Type I, Group VI, Method C, per MIL-I-25135. a. Penetrant, fluorescent, Magnaflux P/N ZL-22A, or equivalent. b. Remover, penetrant, Magnaflux P/N SKC/NF/ZC-7B, or equivalent. c. Developer, penetrant, Magnaflux P/N ZP-9b, or equivalent. d. Black light, portable, Magnaflux P/N ZB-26, or equivalent. 2. Prior to performing the fluorescent penetrant inspection, prepare the airplane by removing the control column assembly from the airplane in accordance with the applicable maintenance document. 3. Gain access for the inspection inside of the airplane in the cockpit, using normal ingress provisions. 4. Prepare the area to be inspected, in accordance with Lockheed Service Manual SMP 515-A or SMP 515-C, Work Card SP-76, Appendix A, for stripping finish around the 1.25-inch diameter hole in the suspect area. WARNING: Paint strippers are dangerous chemicals to humans and aircraft. Every effort must be made to prevent inadvertent contact with any portion of the person or aircraft. Should accidental contact be made, immediately flood the contacted area with water, and, in the case of personal contact, seek immediate medical attention. 5. Perform the fluorescent penetrant inspection in the following manner: NOTE: If the temperature of the part or inspection materials to be used is below 60 degrees F (16 degrees C), preheat and maintain the temperature at or above 60 degrees F (16 degrees C) during application of penetrant materials. The surface may be heated with a hot air heater. a. Pre-clean the area to be inspected with the solvent cleaner (Remover, SKC/NF/ZC-7B) and wipe with a clean, dry, lint free cotton cloth. b. Apply warm air to the inspection area to remove all moisture. c. Apply penetrant to the control column base with penetrant sprayed onto a cotton swab and brushed onto the part. Do not spray the part from the penetrant spray can. Allow a minimum dwell time of 10 minutes before proceeding to the next step. d. Wipe inspection area with a clean, dry, lint free cotton cloth. Wipe remaining penetrant from part with a clean, lint free cotton cloth dampened with the penetrant remover. Check the area to be inspected with the black light to ensure that excess penetrant has been removed prior to applying the developer. e. Spray alight film of developer over the area to be inspected. Allow a minimum dwell time of 5 minutes. f. Using the black light, inspect the suspect area for defects. 6. If defects are confirmed, prior to further flight replace the control column base with a serviceable unit free of cracks. After replacement of the control column base, or if no defects are found, proceed to paragraph C., below (System Securing). C. System Securing: Restore finishes and sealants, reinstall removed components, remove equipment and supplies from the inspection area; rig control system in accordance with applicable maintenance manual; and perform operation checkouts as required in accordance with applicable maintenance instructions. D. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, Atlanta Aircraft Certification Office, FAA, Central Region. E. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. All persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to Lockheed-Georgia Company, 86 South Cobb Drive, Marietta, Georgia 30063. These documents may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or FAA, Central Region, Atlanta Aircraft Certification Office, 1669 Phoenix Parkway, Suite 210, Atlanta, Georgia. This amendment 39-5903 becomes effective June 7, 1988.
93-02-06: 93-02-06 FOKKER: Amendment 39-8488. Docket 92-NM-192-AD. Applicability: Model F27 Mark 100, 200, 300, 400, 500, 600, and 700 series airplanes; certificated in any category. Compliance: Required as indicated, unless accomplished previously. To prevent reduced structural integrity of the wings, accomplish the following: (a) Within 3 years after the effective date of this AD, remove each bolt, part number AN3- 15A or AN3-16A, attaching the truss members to the ribs at wing stations 4800, 5950, 7200, 8350, and 9397, and install two bolts, part number NAS1303, in accordance with Fokker Service Bulletin F27/57-25, Revision 1, dated August 1, 1991. (b) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal MaintenanceInspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113. NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113. (c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (d) The bolt installation shall be done in accordance with Fokker Service Bulletin F27/57-25, Revision 1, dated August 1, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Fokker Aircraft USA, Inc., 1199 North Fairfax Street, Alexandria, Virginia 22314. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (e) This amendment becomes effective on March 12, 1993.
87-25-13: 87-25-13 BRITISH AEROSPACE: Amendment 39-5799. Applies to all Model H.S. 748 series airplanes with Modification 1472 incorporated, but without Modification 7513 incorporated, certificated in any category. Compliance required as indicated, unless previously accomplished. To prevent inability to lower or lock down the nose gear for landing, accomplish the following: A. Prior to the accumulation of 7,000 landings or within the next 90 days after the effective date of this AD, whichever occurs later, replace nose landing gear jack support bracket bearing cap attachment studs, Part Number 2cD13248, in accordance with British Aerospace Alert Service Bulletin A53/53 Revision 1, dated May 1987. B. Replacement studs must, in turn, be replaced prior to accumulation of 7,000 landings. C. Until studs exceeding a life of 7,000 landings have been replaced, nose landing gear jack support structure must be inspected prior to each day's first flight to ensure each stud and bearing cap are secure and correctly fitted in accordance with British Aerospace Alert Service Bulletin A53/53, Revision 1, dated May 1987. D. On assemblies where the bearing caps or studs are found loose, all four bearing cap attachment studs must be replaced before the next flight, in accordance with British Aerospace Alert Service Bulletin A53/53, Revision 1, dated May 1987. E. Incorporation of Modification 7513, as described in British Aerospace Alert Service Bulletin A53/53, Revision 1, dated May 1987, constitutes terminating action for requirements of paragraphs A., B., C., and D., above. F. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety and which has the concurrence of an FAA Principal Maintenance Inspector, may be used when approved by the Manager, Standardization Branch, ANM-113, FAA, Northwest Mountain Region. G. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for the accomplishment of inspections and/or modifications required by this AD. All persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to British Aerospace, Librarian for Service Bulletins, P.O. Box 17414, Dulles International Airport, Washington, D.C. 20041. This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the Seattle Aircraft Certification Office, 9010 East Marginal Way South, Seattle, Washington. This amendment becomes effective January 25, 1988.
2007-10-10 R1: The FAA is revising an existing airworthiness directive (AD), which applies to all Airbus Model A300-600 series airplanes. That AD currently requires revising the Airworthiness Limitations section of the Instructions for Continued Airworthiness to incorporate new limitations for fuel tank systems. This AD clarifies the intended effect of the AD on spare and on-airplane fuel tank system components. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent the potential of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors caused by latent failures, alterations, repairs, or maintenance actions, could result in fuel tank explosions and consequent loss of the airplane. DATES: This AD is effective December 28, 2009. The Director of the Federal Register approved the incorporation by reference of a certain publication listed in the AD as of December 28, 2009. On June 27, 2007 (72 FR 28827, May23, 2007), the Director of the Federal Register approved the incorporation by reference of certain other publications listed in the AD. We must receive any comments on this AD by January 25, 2010.
76-20-03: 76-20-03 PRATT & WHITNEY AIRCRAFT: Amendment 39-2731 as amended by Amendment 39-3070. Applies to all Pratt & Whitney Models JT9D-3A, -7, -7H, -7A, -7AH, -7F and -20 turbofan engines incorporating labyrinth solid land seals P/N's 589733, 589741, 647189 or 729774. I. To prevent possible engine fires due to excessive No. 3 bearing compartment labyrinth seal clearances, accomplish the following: A. Measure the breather air temperature in accordance with Pratt & Whitney Alert Service Bulletin No. 4391, dated February 19, 1975, or later FAA approved revision, or Service Bulletin No. 4703, dated July 1, 1977, or later FAA approved revision, whenever any of the following major engine sections have been or are removed or replaced: 1. Intermediate case. 2. Rear compressor rotor and stator assembly. 3. Rear compressor drive turbine rotor assembly due to: a. Turbine blade root fracture. b. Multiple turbine blade airfoil fracture. c. Failure that causes release of any other rotating part. 4. Front compressor drive turbine rotor assembly due to loss of complete blade. 5. Turbine exhaust case due to loose or missing tailcone. 6. Diffuser case. B. Compare the measured breather air temperature with the limit defined by Pratt & Whitney Curve 4391, for on-ground temperature checks, or Curve 4703, for in-flight temperature checks, and accomplish the following: 1. If the measured breather air temperature is above the limit, remove the engine from service prior to further flight. 2. If the measured breather air temperature is below the limit, accomplish a repetitive temperature check in accordance with the following schedule, or 200 hours after the effective date of this revision, whichever is later: TEMPERATURE MARGIN INSPECTION INTERVAL Less than 5 degrees C 200 hours 5 degrees C but less than 10 degrees C 500 hours 10 degrees C or greater 1000 hours II. To improve the capability of promptly detecting compartment fires, replace main bearing breather tube steel elbow P/N 708237 with magnesium elbow P/N 613265 in accordance with Pratt & Whitney Service Bulletin 4560, dated March 9, 1976, or later FAA approved revision, no later than March 31, 1977. Upon request of the operator, an equivalent method of compliance with the requirements of this AD may be approved by the Chief, Engineering and Manufacturing Branch, Federal Aviation Administration, New England Region. Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Engineering and Manufacturing Branch, New England Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for that operator. The manufacturer's specifications and procedures identified and described in this directive areincorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Pratt & Whitney Aircraft, Division of United Technologies Corporation, 400 Main Street, East Hartford, Connecticut 06108. These documents may also be examined at Federal Aviation Administration, New England Region, 12 New England Executive Park, Burlington, Massachusetts, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at New England Region. Amendment 39-2731 supersedes Amendment 39-2109 (40 FR 8544), AD 75-05-16, as amended by Amendment 39-2137 (40 FR 12773). Amendment 39-2731 became effective 30 days after publication in the Federal Register. This amendment 39-3070 becomes effective November 16, 1977.
94-17-06: This amendment adopts a new airworthiness directive (AD) that applies to GROB Luft und Raumfahrt (Grob) Models G102, G103, G109, and G109B gliders. This action requires inspecting (one-time) the airbrake stops for cracks in the surrounding gelcoat and to ensure that the outer airbrake swivel levers are in contact with the stops during operation, and repairing any gelcoat cracks or any airbrake stops not in contact with the swivel levers. Excessive wear caused the airbrake fence to jam on the upper shell of the wing on one of the affected gliders, resulting in an accident. The actions specified by this AD are intended to prevent airbrake failure caused by jamming of the airbrake fence, which could result in loss of control of the glider.
84-24-51 R1: 84-24-51 R1 THE DeHAVILLAND AIRCRAFT COMPANY OF CANADA, A DIVISION OF BOEING OF CANADA, LTD.: Amendment 39-5030 as amended by Amendment 39-5661. Applies to all Model DHC-7 series airplanes, certificated in any category. Compliance is required as indicated, unless previously accomplished. To ensure correct operation of the airplane during flight in icing conditions, accomplish the following: A. Within 60 days after the effective date of this AD, revise Section 2 of the FAA- approved DHC-7 Airplane Flight Manual (AFM), Revision 29, dated July 26, 1985, as follows, and provide this information to the flight crews: 1. Change the title of paragraph 2.21 to read: "2.21. Flight in Icing Conditions (Visible Moisture and/or Precipitation at Temperatures Below Plus 5 Degrees Centigrade True Outside Air Temperature, or Plus 13 Degrees Centigrade Indicated Outside Air Temperature When Correction Chart Figure 4-4-4 is Not Used)." 2. Change item 8 in paragraph 2.21.1 toread: "8. Ignition switch - Manual (if required)." 3. Add the following note at the end of paragraph 2.21.1: "NOTE: Manual (continuous) engine ignition is an automatic function when AIRFRAME - FAST/SLW switch is at fast or slow position." B. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. C. An alternate means of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, New York Aircraft Certification Office, FAA, New England Region. All persons affected by this directive who have not already received the appropriate service information from the manufacturer may obtain copies upon request to de Havilland Aircraft Company of Canada, A Division of Boeing of Canada, Ltd., Garratt Boulevard, Downsview, Ontario M3K 1Y5, Canada. This information may be examined at FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at FAA, New England Region, New York Aircraft Certification Office, 181 South Franklin Avenue, Room 202, Valley Stream, New York. Amendment 39-5030 became effective April 16, 1985. This amendment, 39-5661, becomes effective August 3, 1987.
80-26-02: 80-26-02 GATES LEARJET: Amendment 39-4015. Applies to 23, 24, 25, 28 and 29 series airplanes certificated in all categories. COMPLIANCE: Required as indicated, unless previously accomplished in accordance with AD 80-22-10. A) Before further flight: 1. Deactivate the pitch function of the FC-110 Automatic Flight Control System (AFCS) or Automatic Flight Control Stability System (AFC/SS), as indicated below, by pulling the AFCS Pitch DC Circuit Breaker to the off position, banding it to prevent use of this function and checking to assure this function is the only deactivated circuit or control: SERIES SERIAL NUMBERS LOCATION 23 003 thru 014 015 thru 099 Pilot's Switch Panel Pilot's Sub Panel 24 100 thru 139 (except 131, 132, & 134) 131, 132, & 134 140 thru 229 230 and up Pilot's Sub Panel Pilot's circuit breaker panel Autopilot computer rack (under pilot's seat) Pilot's circuit breaker panel 25 003 thru 069 (except 032) 032070 and up Autopilot computer rack (under pilot's seat) Pilot's Sub Panel Pilot's circuit breaker panel 28 001 and up Pilot's circuit breaker panel 29 001 and up Pilot's circuit breaker panel 2. Install a locally fabricated placard on or near the autopilot control head in clear view of the crew, using letters at least 3/32 inch high, which reads: AUTOPILOT PITCH AXIS INOPERATIVE OBSERVE APPROPRIATE AFM AIRSPEED LIMITATIONS FOR INOPERATIVE AUTOPILOT and operate the airplane in accordance with this placard. 3. Insert in the appropriate section of the existing Airplane Flight Manual (AFM) the FAA approved temporary Airplane Flight Manual Change dated October 22, 1980, pertaining to emergency procedures for pitch axis malfunction. B) On or before April 1, 1981, accomplish all of the following at a Gates Learjet authorized service center holding appropriate FAA repair station ratings. 1. Visually inspect the elevator control system to assurethat Pitch Axis Servo (D.C. Torquer), P/N 6600163-( ) is installed. a) If installed, modify the airplane by incorporating autopilot pitch trim monitor test switch in accordance with Gates Learjet Airplane Modification Kit AMK 80- 16B, Change 2. b) If not installed, modify the airplane by replacing the pitch servo actuator and capstan and incorporating autopilot pitch trim monitor test switch in accordance with Gates Learjet Airplane Modification Kits AMK 80-3, Change 4, and AMK 80-16B, Change 2, respectively. 2. Insert in the appropriate sections of the existing Airplane Flight Manual (AFM) the FAA approved temporary Airplane Flight Manual changes dated October 22, 1980, for autopilot trim monitor. C) When paragraph B of this AD has been accomplished, the requirements of paragraphs A)1. and 2. of this AD are no longer applicable. D) Airplanes may be flown in accordance with FAR 21.197 to a location where the requirements of this AD can be accomplished provided the autopilot is not operative during that flight. E) Any equivalent method of compliance with this AD must be approved by the Chief, Aircraft Certification Program, FAA, Central Region, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. This AD supersedes AD 80-22-10. This amendment becomes effective on January 15, 1981, to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated December 11, 1980, and is identified as AD 80-26-02.
2009-25-13: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: There have been 3 reported occurrences of uncontrolled excessive heat from the left hand baggage bay sidewall heater, [part number] P/N 3436-06-1/0, that resulted in the affected sidewall heater panels sustaining heat discoloration and/or scorching of the liner material. The affected sidewall heater is equipped with a thermostat to regulate heating. These reported occurrences are the subject of further investigation. As a preventive measure, until such time as the cause of the occurrences have been determined, deactivation of the left hand baggage bay heater is necessary to avoid the potential for uncontrolled excessive heat by the heater panel, and on the baggage bay compartment, that could lead to flammability issues. * * * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
2009-22-13: We are adopting a new airworthiness directive (AD) for certain Boeing Model 767-200, -300, -300F, and -400ER series airplanes. This AD requires an inspection to determine if certain motor operated valve actuators for the fuel tanks are installed, and related investigative and corrective actions if necessary. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent an ignition source inside the fuel tanks, which, in combination with flammable fuel vapors, could result in a fuel tank explosion and consequent loss of the airplane.
92-01-06: 92-01-06 BOEING: Amendment 39-8130. Docket No. 91-NM-153-AD. \n\n\tApplicability: Model 757 series airplanes, as listed in Boeing Service Letter 757-SL-28- 8-A, dated June 5, 1991, certificated in any category. \n\n\tCompliance: Required within the next 3,000 hours time-in-service after the effective date of this AD, unless previously accomplished. \n\n\tTo assure that the fuel crossfeed valve and fuel shutoff valves close fully, accomplish the following: \n\n\t(a)\tPerform a visual inspection or a functional test of the fuel crossfeed valve and fuel shutoff valves to determine if the valves are correctly installed, in accordance with Boeing Service Letter 757-SL-28-8-A, dated June 5, 1991. If an incorrectly installed valve is found, prior to further flight, remove and re-install it correctly, in accordance with the service letter.\n \n\t(b)\tWithin 30 days after accomplishing the inspection or functional test required by paragraph (a) of this AD, submit a report of the inspection ortest findings from which it is determined that the fuel crossfeed valve or fuel shutoff valves were incorrectly installed, to: Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region, 1601 Lind Avenue SW., Renton, Washington 98055; rapid fax: (206) 227-1181; telex 756366. Information collection requirements contained in this regulation have been approved by the Office of Management and Budget (OMB) under the provision of the Paperwork Reduction Act of 1980 (P.L. 96-511) and have been assigned OMB Control Number 2120-0056. \n\n\t(c)\tAn alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Seattle Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Seattle ACO. \n\n\t(d)\tSpecial flight permitsmay be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base in order to comply with the requirements of this AD. \n\n\t(e)\tThe inspection/test requirements of this AD shall be done in accordance with Boeing Alert Service Letter 757-SL-28-8-A, dated June 5, 1991. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Boeing Commercial Airplane Group, P.O. Box 3707, Seattle, Washington 98124. Copies may be inspected at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, Washington, or at the Office of the Federal Register, 1100 L Street NW., Room 8401, Washington, D.C. \n\n\t(f)\t This amendment (39-8130, AD 92-01-06) becomes effective on February 17, 1992.
77-12-07: 77-12-07 BOEING: Amendment 39-2924. Applies to Model 747 airplanes certificated in all categories and equipped with fuel valve actuator assemblies, Boeing P/N 60B92406-27 (ITT General Controls P/N MA11A1173, serial numbers H78442 through H98252 and J00047 through J34689) or Boeing P/N 60B92406-28 (ITT General Controls P/N MA11A1174, serial numbers H95059 through H95074 and J04218 through J37314) installed or manufactured between April 30, 1975, and April 1, 1977. Compliance required as indicated. \n\tTo prevent loss of control of fuel shutoff, crossfeed and required fuel management capability, accomplish the following: \n\tA.\tWithin 145 days or 1300 hours time in service after the effective date of this Airworthiness Directive whichever occurs first, remove fuel valve actuator assembly, Boeing P/N 60B92406-27 (ITT General Controls P/N MA11A1173) or Boeing P/N 60B92406-28 (ITT General Controls P/N MA11A1174) from the No. 1, 2, 3 and 4 engine fuel shutoff valves, AND \n\t\t(1)\treplace with actuator assembly Boeing P/N 60B92406-32 (ITT General Controls P/N MA11A1173-1) or Boeing P/N 60B92406-33 (ITT General Controls P/N MA11A1174-1): OR \n\t\t(2)\treplace with an actuator assembly which has been inspected and reworked to incorporate a properly heat treated armature shaft in accordance with ITT General Controls Service Bulletins MA11A1173-24-01 or MA11A1174-24-01, both dated March 30, 1977, and 107728A101-24-01, dated June 10, 1977, or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA Northwest Region. \n\tB.\tWithin 1 year or 3200 hours time in service after the effective date of this Airworthiness Directive whichever occurs first, replace or rework the actuator assemblies Boeing P/N 60B92406-27 (ITT General Controls P/N MA11A1173) or Boeing P/N 60B92406-28 (ITT General Controls P/N MA11A1174) which are installed on the No. 1, 2, 3 and 4 crossfeed valves, the No. 1 and 4 reserve tank gravity transfer valves, the center wing tank fuel jettison valves and, if installed, the No. 2 and 3 reserve tank gravity transfer valves in accordance with A.(1) or A.(2) above. \n\tThe manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to U.S.C. 552(a)(1). \n\tAll persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to ITT General Controls, Aerospace Products, 1200 South Flower Street, Burbank, California 91502. These documents may also be examined at the Engineering and Manufacturing Branch, FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. \n\n\tThis amendment becomes effective June 20, 1977.
85-18-05 R2: 85-18-05 R2 SIKORSKY AIRCRAFT: Amendment 39-5129 as amended by Amendment 39-5525 is further amended by Amendment 39-6098. Applicability: Model S-61L, S-61N, S-61NM, and S-61R series helicopters, certificated in all categories, and S-61A (S/N's 61083, 61087, 61094, and 61161) and S-61V (S/N 61271) helicopters, certificated in the restricted category, which are engaged in more than six external cargo lifts per flight hour under Part 133, Class B, Rotorcraft external load combination operations. Compliance: Required as indicated (unless already accomplished). To prevent operation with a main rotor spar crack and possible loss of the helicopter, accomplish the following: (a) Within the next 25 hours time in service after the effective date of this AD, unless already accomplished, remove main rotor blades from the rotorcraft that are not approved for use in Part 133 (Class B, Rotorcraft-external load combination operations), and replace with approved blades. The approved main rotor blades are as follows: (1) The following blades are approved for Model S-61L, transport category helicopters operating up to a combined vehicle and cargo gross weight of 22,000 lbs., provided the main rotor blades have been altered and maintained in accordance with Service Bulletin (SB) 61B15-6, Rev. P, or later FAA-approved revisions, excluding Section 2, Part II: (i) P/N's S6115-20501-041 and -042. (ii) P/N's S6115-20601-042, -047, and -048. (iii) P/N's 61170-20201-055, -056, -058, -059, -060, -061, -062, -065, and -067. (iv) P/N's S6117-20101-041, -046, -050, -051, -054, -055, -056, -057, and 058. (2) The following blades are approved for Model S-61N, transport category helicopters operating up to a combined vehicle and cargo gross weight of 22,000 lbs., or Model S-61NM, transport category helicopters operating up to a combined vehicle and cargo gross weight of 20,500 lbs., provided the main rotor blades have been altered and maintained in accordance with SB No. 61B15-6, Rev. P, or later FAA-approved revisions, excluding Section 2, Part II: (i) P/N's S6115-20501-041, and -042. (ii) P/N's S6115-20601-041, -042, -045, -046, -047, and -048. (iii) P/N's S6188-15001-041 and -045. (iv) P/N's 61170-20201-054, -055, -056, 058, -059, 060, -061, -062, -065, -067. (v) P/N's S6117-20101-041, -046, -050, -051, -054, -055, -056, -057, and -058. (3) P/N's 61170-20201-062 blades are approved for the Model S-61A (S/N's 61083 and 61094), restricted category helicopters, operating up to a combined vehicle and cargo gross weight of 22,000 lbs. (4) P/N's S6115-20201-2 and -3 blades are approved for the Model S-61A (S/N's 61087 and 61161), restricted category helicopter, operating up to a combined vehicle and cargo gross weight of 19,000 lbs. (5) P/N 61170-20201-060 blades are approved for the Model S-61V (S/N 61271), restricted category helicopter, operating up to a combined vehicle and cargo gross weight of 19,100 lbs. (6) The following blades are approved for Model S-61R transport category helicopters operating up to a combined aircraft and cargo gross weight of 19,500 pounds: (i) P/N's S6115-20501-041 and -042. (ii) P/N's S6115-20601-042, and -045 through -048. (iii) P/N's S6117-20101-041, -050, -051, -054, -056, -057, and -058. (iv) P/N's 61170-20201-055, -056, -058 through -062, -064, -065, and -067. (b) Within the next 1 1/2 hours time in service after the effective date of this AD, unless already accomplished, inspect main rotor blades equipped with approved visual blade pressure indicators (VBIM) but not equipped with an in-cockpit blade inspection system (CBIM) in accordance with paragraph (c). After the initial inspection, conduct further inspections in accordance with paragraph (c) prior to the first flight of each day and conduct subsequent visual inspections of the VBIM indicators in accordance with Section 2, Part IV, paragraph la of Sikorsky Service Bulletin No. 61B15-6, Revision P, or later FAA-approved revisions, at intervals not to exceed 1 1/2 hours time in service from the last inspection. (c) Inspect the VBIM indicators of the main rotor blades in accordance with procedures set forth in Section 2, Part IV, of Sikorsky SB No. 61B15-6 Rev. P, or later FAA-approved revisions. (1) Conduct visual inspection of blade-mounted VBIM indicators from the transmission work platform of the helicopter or equivalent to ensure that an accurate visual check is conducted. (2) The visual inspection of blade-mounted VBIM indicators shall be conducted by either an individual who holds a pilot certificate with appropriate rating, or a mechanic certificate with airframe rating, or by an appropriately certificated maintenance entity. The person performing this inspection or check shall make entries of the results in the aircraft maintenance record including a description and date of the inspection and the name of the individual performing the inspection along with the certificate number, kind of certificate, and signature. (d) For helicopters equipped with in-cockpit CBIM (reference Sikorsky SB No. 61B15-20D). (1) Prior to the first flight of the day, after the effective date of this AD, unless already accomplished, and every 8 hours time in service thereafter. (i) Visually inspect the main rotor blade VBIM pressure indicators in accordance with paragraph (c). (ii) Test the VBIM pressure indicators and the in-cockpit CBIM transducers in accordance with the procedures set forth in Section 2, Part IV, of Sikorsky SB No. 61B15-6, Rev. P, or later FAA-approved revisions. (2) Check the in-cockpit blade inspection system electrical circuit and CBIM warning light in flight by activating the (cockpit) BIM test switch located on the left overhead quarter panel at least once each (1) hour time in service during flight operations in accordance with the Rotorcraft Flight Manual (RFM). (i) If the (cockpit) BIM warning light illuminates, continue operations in a normal manner. (ii) If the (cockpit) BIM warning light does not illuminate, immediately check the BIM circuit breaker and reset if tripped. (A) Repeat check of (cockpit) BIM test switch to verify if warning light illuminates. Continue with normal operations if BIM warning light functions properly. (B) If the (cockpit) BIM warning light fails to illuminate, discontinue external load operations and land as soon as practical. Investigate and correct malfunction prior to further flight. (3) If the (cockpit) BIM warning light illuminates during flight, discontinue external load operations and follow the appropriate emergency flight procedures in Part I, Section III, of the SA 4045-30 (S-61L) SA4045-100 (S-61L), or SA4045-82 (S-61N) RFM's. NOTE: For Model S-61 helicopters not engaged in Part 133 external load operations, AD 74-20-07, Rev. 5, main rotor blade inspection requirements are applicable. (e) Each blade with any black or red indication visible in the blade VBIM pressure indicator (or whose transducer activates the cockpit BIM warning light) is restricted from further flight until the cause of the indication is determined and corrected in accordance with procedures given in Sikorsky SB 61B15-6, Rev. P, or later FAA-approved revisions. (f) Alternate inspections, repairs, modifications, or other means of compliance which provide an equivalent level of safety may be approved by the Manager, Boston Aircraft Certification Office, FAA, New England Region, 12 New England Executive Park, Burlington, Massachusetts 01803. (g) Rotorcraft may be flown in accordance with the provisions of FAR Sections 21.197 and 21.199 to a base where the AD can be accomplished, except when a VBIM or CBIM indication exists. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552 (a) (1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to Sikorsky Aircraft, Division of United Technologies, North Main Street, Stratford, Connecticut 06601, Attn: S-61 Commercial Product Support Department. These documents also may be examined at the Office of the Regional Counsel, Southwest Region, FAA, Bldg. 3B, 4400 Blue Mound Road, Fort Worth, Texas 76106. This amendment revises Amendment 39-5129 (50 FR 38506; September 23, 1985), AD 85-18-05, as amended by Amendment 39-5525 (52 FR 8582; March 19, 1987), AD 85-18-05 R1 which was effective on April 13, 1987. This amendment (39-6098, AD 85-18-05 R2) becomes effective February 8, 1989.
2009-25-04: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) originated by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: An operator of A330 aeroplane fitted with Rolls-Royce (RR) Trent 772 B engines experienced an engine1 uncontained multiple turbine blade failure. Investigations have shown that High Pressure/ Intermediate Pressure (HP/IP) oil vent tubes are prone to be affected by carbon deposit or to be damaged by their outer heat shields leading to a fire inside or outside the vent tube and resulting into IP Turbine (IPT) disc drive arm fracture and thus IPT disc overspeed. If not corrected, IPT disc overspeed could lead to an uncontained engine failure, i.e. multiple turbine blade failure or HP/IP turbine disc burst, which would constitute an unsafe condition. ** * * * This AD requires actions that are intended to address the unsafe condition described in the MCAI.
93-23-01: 93-23-01 NORDSKOG INDUSTRIES, INC.: Amendment 39-8735. Docket 93-NM-181-AD. \n\n\tApplicability: Nordskog water heaters and coffee makers, as listed in Nordskog Industries, Inc., Service Bulletin SB-93-34, dated October 21, 1993; as installed in, but not limited to Boeing Model 727, 737, 747, 757, and 767 series airplanes; McDonnell Douglas Model DC-9, DC-9-80, and DC-10 series airplanes, and MD-11 airplanes; Lockheed Model L-1011 series airplanes; Airbus Industrie Model A300, A310, and A320 series airplanes; Gulfstream Model G-1159 series airplanes and Model G-IV airplanes; de Havilland, Inc., Model DHC-8 series airplanes; Dassault-Aviation Model Mystere-Falcon 50 and 900 series airplanes; Canadair Model CL-600-1A11 (CL-600), CL-600-2A12 (CL-601), and CL-600-2B16 (CL-601-3A and -3R) and CL-600-2B19 series airplanes; and Fokker Model F27 and F28 series airplanes; certificated in any category. \n\n\tCompliance: Required as indicated, unless accomplished previously. \n\n\tTo preventexplosions of galley water heaters and coffee makers and subsequent injuries to passengers or cabin crew members, accomplish the following: \n\n\t(a)\tWithin 30 days after the effective date of this AD, perform a one-time inspection to determine whether a NUPRO pressure relief valve having part number (P/N) SS-2C4-65 has been installed, in accordance with Nordskog Industries, Inc., Service Bulletin SB-93-34, dated October 21, 1993. If any NUPRO pressure relief valve having P/N SS-2C4-65 has been installed, prior to further flight, accomplish either paragraph (a)(1) or (a)(2) of this AD. \n\n\t\t(1)\tRemove the NUPRO pressure relief valve having P/N SS-2C4-65 and install a new, improved NUPRO pressure relief valve having P/N SS-CHF2-65, in accordance with the service bulletin. Or \n\n\t\t(2)\tDeactivate any Nordskog water heater or coffee maker listed in the service bulletin on which a NUPRO pressure relief valve having P/N SS-2C4-65 has been installed, and install a placard stating, "Not tobe used." \n\n\t(b)\tAs of the effective date of this AD, no person shall install a NUPRO pressure relief valve having P/N SS-2C4-65 on any airplane. \n\n\t(c)\tAn alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Los Angeles ACO. \n\n\tNOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Los Angeles ACO. \n\n\t(d)\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. \n\n\t(e)\tThe inspection and replacement shall be done in accordance with Nordskog Industries, Inc., Service Bulletin SB-93-34, dated October 21, 1993. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Nordskog Industries, Inc., 16000 Strathern Street, Van Nuys, California 91406. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the FAA, Los Angeles ACO, 3229 E. Spring Street, Long Beach, California 90806-2425; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. \n\n\t(f)\tThis amendment becomes effective on December 7, 1993.
2007-21-14 R1: The FAA is revising an existing airworthiness directive (AD), which applies to all Airbus Model A310 airplanes. That AD currently requires revising the Airworthiness Limitations Section of the Instructions for Continued Airworthiness to incorporate new limitations for fuel tank systems. This AD clarifies the intended effect of the AD on spare and on-airplane fuel tank system components. This AD results from fuel system reviews conducted by the manufacturer. We are issuing this AD to prevent the potential of ignition sources inside fuel tanks, which, in combination with flammable fuel vapors caused by latent failures, alterations, repairs, or maintenance actions, could result in fuel tank explosions and consequent loss of the airplane.
91-12-17: 91-12-17 McDONNELL DOUGLAS: Amendment 39-7029. Docket No. 90-NM-280-AD. \n\n\tApplicability: All McDonnell Douglas Model DC-9-15F, -32F, -33F, -34F, and C-9 (military) series airplanes, certificated in any category. \n\n\tCompliance: Required as indicated, unless previously accomplished. \n\n\tTo prevent inadvertent opening of the forward upper cargo door in flight, a condition which could result in loss of pressurization and reduced controllability of the aircraft, accomplish the following: \n\n\tA.\tWithin four months after the effective date of this AD, and thereafter at intervals not to exceed one year, perform magnetic particle inspections on the cargo door latch spool fitting attach bolts or replace the non-Inconel cargo door latch spool fitting attach bolts with new bolts, in accordance with the Accomplishment Instructions for Phase 2 of McDonnell Douglas DC-9 Alert Service Bulletin A52-174, dated August 7, 1990, or Revision 1, dated December 14, 1990. \n\n\t\t1.\tIf a bolt doesnot pass the magnetic particle inspection, prior to further flight, replace it with a new bolt and seal in accordance with the service bulletin. \n\n\t\t2.\tIf a bolt passes the magnetic particle inspection, prior to further flight, reinstall the bolt and seal in accordance with the service bulletin. \n\n\tB.\tThe inspections required by paragraph A. of this AD are not required for Inconel bolts, part numbers RA21026-7-28, 77711-7-28, and 3D0031-7-28. \n\n\tC.\tWithin two years after the effective date of this AD, replace all non-Inconel cargo door latch spool fitting attach bolts with Inconel bolts, part numbers RA21026-7-28, 77711-7-28, or 3D0031-7-28, in accordance with the Accomplishment Instructions of Phase 3 of McDonnell Douglas DC-9 Alert Service Bulletin A52-174, dated August 7, 1990; or Phase 4 of McDonnell Douglas DC-9 Alert Service Bulletin A52-174, Revision 1, dated December 14, 1990. Installation of Inconel bolts constitutes terminating action for the requirements of paragraph A. of this AD. \n\n\tD.\tSpecial flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes unpressurized to a base for the accomplishment of the requirements of this AD. \n\n\tE.\tAn alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Los Angeles Aircraft Certification Office (ACO), FAA, Transport Airplane Directorate. \n\n\tNOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment and then send it to the Manager, Los Angeles ACO. \n\n\tAll persons affected by this directive who have not already received the appropriate service documents from the manufacturer may obtain copies upon request to McDonnell Douglas Corporation, P.O. Box 1771, Long Beach, California 90846-0001, Attention: Business Unit Manager, Technical Publications, C1-HCW (54-60). These documents may be examined at the FAA, Northwest Mountain Region, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or the Los Angeles Aircraft Certification Office, 3229 East Spring Street, Long Beach, California. \n\n\tThis amendment (39-7029, AD 91-12-17) becomes effective on July 15, 1991.
2009-25-02: We are adopting a new airworthiness directive (AD) for certain Twin Commander Aircraft LLC Models 690, 690A, and 690B airplanes. This AD requires you to inspect between the surface of the left-hand (LH) and right-hand (RH) upper wing skins and the engine mount beam support straps for any signs of corrosion, replace the upper steel straps with parts of improved design, and modify both wings. This AD results from reports that corrosion was found between the mating surfaces of the wing upper skin surface and the engine mount beam support straps. We are issuing this AD to detect and correct corrosion on the engine mount beam support straps and the upper wing skins, which could result in failure of the engine mount beam support straps. This failure could lead to loss of the engine and possible loss of control of the airplane.